Aircraft

ABSTRACT

An aircraft comprises at least first and second gas turbine engines arranged in a line extending generally normally to an aircraft longitudinal axis (A), each engine comprising at least one compressor or turbine rotor disc defining a respective rotational plane (D 32 -D 42 ). The rotational plane (D 32 -D 42 ) of at least one of the rotors of at least one of the engines is angled relative to the aircraft longitudinal line (A) such that a burst disc plane of the respective engine is nonintersecting with another engine.

PRIORITY

This is a Continuation of application Ser. No. 14/976,414 filed Dec. 21,2015, which claims priority to British Patent Application No. 1500996.2filed Jan. 21, 2015. The disclosure of the prior applications is herebyincorporated by reference herein in their entirety.

FIELD OF THE INVENTION

The present invention relates to an aircraft, particularly though notexclusively, to a blended wing body aircraft having multiple engineslocated at a trailing edge of the aircraft.

BACKGROUND TO THE INVENTION

Blended wing body aircraft are known in which the fuselage and wings areintegrated such that there is no clear dividing line between the wingsand fuselage. The body may therefore contribute to lift, increasing theefficiency of the aircraft. In a “flying wing” design, the fuselage isomitted, with the pilot, passengers and engines provided within thewing.

One known blended wing body study is described in Liebeck, R. H. “Designof the Blended Wing Body Subsonic Transport.” AIAA Journal of Aircraft,Volume 41, Issue 1, January-February 2004, pp. 10-25. Blended wing bodyand flying wing configurations are also known in which the engines aremounted within the wing (such as the Northrop B2). Trade studies haveshown that such a design is most suitable for relatively large, longrange aircraft.

Gas turbine engines are highly efficient, and have high thrust to weightratios. However, some failure modes of gas turbine engines can causeextensive damage to safety critical aircraft components such as otherengines, fuel tanks, hydraulic and electrical control runs, and aircraftpropulsors in the case of a distributed propulsion aircraft. Forexample, the uncontained rupture of a compressor or turbine rotor disccan, in some cases, lead to the destruction of adjacent engines, firesin fuel tanks, or severance of hydraulic or electrical control runs,which may lead to loss of control. Consequently, in most conventionalaircraft, the engines are spaced far apart (either by locating enginesspaced on the wing, or by locating one or more engines in the tail ofthe aircraft) and physically distant from safety critical aircraftcomponents, thereby reducing the risks of damage to other safetycritical aircraft components in the event of an uncontained disc ruptureevent. Alternatively, the engines may be surrounded by relatively heavyand expensive containment structures to prevent uncontained failuresfrom occurring.

However, it is desirable to locate both the wings and engines close tothe centre of mass of the aircraft, since the engines represent a largeproportion of the mass of the aircraft, and locating the wings (and sothe centre of lift) close to the centre of mass reduces trim drag whichwould otherwise be necessary to cancel the torque produced in flight bythe centre of lift/centre of mass mismatch. If the engines are installedwithin the fuselage, this leads to engines close to one another, as wellas close to fuel tanks in the wings, and hydraulic and/or electricalcontrol runs which pass from the cockpit to the aft of the aircraft.Such an arrangement, while aerodynamically and structurally efficient,leads to unacceptable aircraft risks in the event of an uncontained discfailure.

The present invention describes an aircraft which seeks to overcome someor all of the above problems.

SUMMARY OF THE INVENTION

According to a first aspect of the present invention, there is providedan aircraft having a longitudinal axis and comprising:

at least one gas turbine engine having a longitudinal axis andcomprising at least one compressor or turbine rotor disc defining arespective rotational plane normal to the longitudinal axis, wherein theengine longitudinal axis is angled relative to the aircraft longitudinalaxis such that a respective rotational plane of the respective engine isnonintersecting with one or more safety critical aircraft component; and

at least one duct arrangement comprising an intake duct leading from anair inlet to an intake of a respective engine, and an exhaust ductleading from an exhaust of the respective engine to an exhaust outlet,wherein the intake duct transitions from a direction generally parallelto the aircraft longitudinal axis at the air inlet, to a directiongenerally parallel to the respective engine longitudinal axis at theintake, and the outlet duct transitions from a direction generallyparallel to the respective engine longitudinal axis at the exhaust ofthe respective engine, to a direction generally parallel to the aircraftlongitudinal axis at the exhaust outlet.

Advantageously, the design freedom for aircraft engine placement isincreased, since the engine can be placed at a longitudinal positionclose to safety critical aircraft components, without riskingdestruction of these components in the event of a disc burst event. Theintake arrangement permits such an arrangement without significantreduction in engine efficiency, since inlet pressure is recovered by theintake arrangement, and the direction of thrust is redirected to theaircraft longitudinal line by the exhaust arrangement. This arrangementhas been found by the inventor to permit aircraft designs havingsignificantly lower weight, and/or higher aerodynamic efficiencycompared to prior arrangements.

The one or more safety critical aircraft components may comprise one ormore of a further gas turbine engine, a fuel tank, a hydraulic controlrun, an electrical control run, a passenger or crew compartment, and anaircraft propulsor.

The aircraft may comprise two or more gas turbine engines and associatedduct arrangements, which may be arranged in a line extending generallynormally to an aircraft longitudinal axis in the horizontal plane.

The one or more gas turbine engines may be arranged in a linesubstantially coincident with an aircraft centre of mass, and may bearranged in a line substantially coincident with an aircraft centre oflift. Advantageously, each of the aircraft centre of lift, engines andcentre of mass can be substantially coincident, without risking damageto safety critical aircraft components in the event of a disc burstevent.

The one or more gas turbine engines may be offset relative to theaircraft longitudinal axis in the horizontal plane.

The aircraft may comprise one of a blended wing body aircraft and aflying wing aircraft.

The engines may be located at a trailing edge of the fuselage of theaircraft, and may be mounted at a point of transition between thefuselage and wings. The engines may be located within the aircraft wingor fuselage.

The air inlet of the or each intake duct may be located at a leadingedge of the point of transition between the fuselage and wings.Advantageously, the intake is located at a stagnation point of airpassing over the aircraft, thereby increasing the static pressure of airentering the engine.

The aircraft may comprise at least two redundant control runs for atleast one aircraft control surface, wherein one control run extendsalong a first side of an engine, and a further control run extends alonga second side of the respective engine Advantageously, an uncontaineddisc failure (which will normally only produce a small number ofrelatively large fragments) is unlikely to sever both control runs inview of the large separation between redundant control runs.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 shows a schematic plan view of an aircraft in accordance with thepresent disclosure; and

FIG. 2 shows a cross sectional schematic view of a gas turbine enginesuitable for the aircraft of FIG. 1.

DETAILED DESCRIPTION

FIGS. 1 and 2 show an aircraft 2 in accordance with the presentdisclosure. The aircraft is in the form of a blended wing body (BWB)business jet aircraft of a size suitable for carrying up to 10 people.

The aircraft 2 includes a body 4 configured to carry passengers within aforward cabin section 5, wings 6 blended with the aircraft body, andengines 10 a, 10 b located on either side of the body 4. The cabinsection 5 is oval when viewed from above. Such a configuration providesa structurally efficient pressure vessel for the passengers and crew,while providing a large internal volume. The cabin section 5 has a loweraspect ratio (i.e. ratio of length to width) than a conventionalpassenger cabin—this is thought to increase the productivity of thespace, since passengers can face one another. The aft fuselage volumecan be utilised for non-passenger cabin aircraft internal volumerequirements, such as avionics, unpressurised or pressurised cargo orluggage space, and fuel tanks 58. The aircraft wings 6 are relativelyhigh aspect ratio, and provide a portion of the lift for the aircraft 2in flight, the remainder of the lift being provided by the body 4, whichis of a “lifting body” configuration, being aerofoil shaped, andsmoothly transitioning to the wings 6 at a transitioning portion 8. Thebody 4 defines an aircraft longitudinal axis A which extends from a nose7 of the aircraft to a tail 9. A lateral axis B extends through a centreof mass M, normally to the aircraft longitudinal axis A in thehorizontal plane. In the described example, the aircraft is a “businessjet” type aircraft configured to carry approximately 10 or fewerpassengers. The described aircraft has a wing span of approximately 40m, a length of approximately 33 m, and a fuselage height ofapproximately 3.3 m. The oval passenger cabin 5 has an internal lengthof approximately 9 m, and an internal width of approximately 7 m.

The engines 10 a, 10 b are located at a trailing edge of an outerportion of the body 4 at the transitioning portion 8, either side of theaircraft longitudinal axis A, and so the engines 10 a, 10 b are offsetfrom the aircraft longitudinal centre line A. The engines 10 a, 10 b areeach in the form of a gas turbine 10. The engine 10 is shown in FIG. 2and comprises, in axial flow series, an intake 11, fan 12, a bypass duct13, an intermediate pressure compressor 14, a high pressure compressor16, a combustor 18, a high pressure turbine 20, an intermediate pressureturbine 22, a low pressure turbine 24 and an exhaust nozzle 25. The fan12, compressors 14, 16 and turbines 20, 22, 24 all rotate about themajor axis of the gas turbine engine 10 and so define the axialdirection of gas turbine engine. By positioning the engines 10 withinthe relatively thick outer portion of the body, sufficient space isprovided to install an engine 10 having a relatively large fan diameter,and thus high bypass ratio. Such engines are generally more efficientthan low bypass ratio engines at high subsonic speeds. On the otherhand, this area is too thin to provide sufficient headroom forpassengers, so would be otherwise wasted.

Air is drawn in through the intake by the fan 12 where it isaccelerated. A significant portion of the airflow is discharged throughthe bypass duct 13 generating a corresponding portion of the engine 10thrust. The remainder is drawn through the intermediate pressurecompressor 14 into what is termed the core of the engine 10 where theair is compressed. A further stage of compression takes place in thehigh pressure compressor 16 before the air is mixed with fuel and burnedin the combustor 18. The resulting hot working fluid is dischargedthrough the high pressure turbine 20, the intermediate pressure turbine22 and the low pressure turbine 24 in series where work is extractedfrom the working fluid. The work extracted drives the intake fan 12, theintermediate pressure compressor 14 and the high pressure compressor 16via shafts 26, 28, 30. The working fluid, which has reduced in pressureand temperature, is then expelled through the exhaust nozzle 25 andgenerates the remaining portion of the engine 10 thrust.

The fan 12, compressors 14, 16 and turbines 20, 22, 24 each comprise aplurality of blades affixed to respective discs 32-42. Each of thesediscs 32-42 rotate at a high speed. In the event of failure of one ofthese discs, fragments of the respective failed disc 32-42 may not becontained within the engine 10, and may therefore exit the engine 10 athigh velocity, which may cause damage to adjacent components. Thistherefore imposes a minimum spacing requirement for the engines 10, andmay necessitate relatively heavy, bulky and expensive containmentsystems for the engines 10.

Referring again to FIG. 1, the engines 10 a, 10 b have respectivelongitudinal axes C_(a), C_(b), which extend from the respective engineintake 11 to the respective exhaust nozzle 25. Each disc 32-42 rotatesabout the respective engine longitudinal axis C, and so defines arespective rotational plane D₃₂-D₄₂ extending normally to the enginelongitudinal axis C. FIG. 2 shows the rotational plane D₃₈ for the highpressure turbine rotor 38 of each engine 10 _(a), 10 _(b). Therespective rotational plane D₃₂-D₄₂ for each disc 32-42 represents thepotential paths that fragments from the discs 32-42 may take in theevent of an uncontained disc failure, and so represents a “burst discplane”.

The engines 10 a, 10 b are angled such that their longitudinal axesC_(a), C_(b), are non-parallel with the aircraft longitudinal axis A inthe horizontal plane, i.e. the engines 10 a, 10 b are angled inwardlysuch that the respective intakes 11 are canted inward toward theaircraft nose 7. The engines 10 a, 10 b are canted inward to an extentsuch that at least some of the rotational planes D₃₂-D₄₂ of the engines10 a, 10 b is nonintersecting with the other engine 10 a, 10 b, i.e. anotional line extending from one or more rotating components of theengine in a direction normal to the axis of rotation of the rotatingcomponents does not extend through the other engine 10 a, 10 b.Consequently, in the event of an uncontained disc failure, discfragments from the engines 10 a, 10 b will not strike the other engine.In particular, it is desirable that the plane D₃₈ of the high energy,otherwise uncontained HP turbine disc 38 is nonintersecting with theother engine 10 a, 10 b. Consequently, safety critical aircraftcomponents in the form of the engines 10 a, 10 b do not intersect therotational planes D₃₂-D₄₂ of the other engines.

Fuel tanks are also regarded as safety critical aircraft components.Consequently, the fuel tank 58 is also located away from the rotationalplanes D₃₂-D₄₂ of the engines 10 a, 10 b, such that the rotationalplanes D₃₂-D₄₂ is nonintersecting with the fuel tank 58. Consequently, adisc burst event will not damage fuel tanks, which may otherwise lead toa fire. Further fuel tanks (not shown) can also be provided in the wings5 between the rotational planes D₃₂-D₄₂ of the engines 10 a, 10 bwithout encountering a risk of damage in the event of a disc rupture.The cabin 5 is also located forward of the rotational planes D₃₂-D₄₂ ofthe engines 10 a, 10 b, and so fragments from a disc burst event willnot enter the passenger cabin.

Since the engines 10 a, 10 b are angled such that their longitudinalaxes C_(a), C_(b) do not correspond to the aircraft longitudinal axis A,any thrust generated by the engines 10 a, 10 b would be directed awayfrom the direction of travel of the aircraft, leading to inefficiency.Consequently, the engines 10 a, 10 b are each located within a ductarrangement in the form of “S-ducts”.

Referring again to FIG. 1, an intake duct 44 is provided for each engine10 a, 10 b. The respective intake ducts 44 extend from a leading edge ofthe transition region 8 of the fuselage 4 to the engine intake 11. Theintake duct 44 transitions from a direction generally parallel to theaircraft longitudinal axis A at an inlet to the intake duct, to adirection generally parallel to the respective engine rotational axis Cat the engine inlet 11. The inlet duct 44 may transition smoothlybetween these directions, to avoid flow separation from the walls of theintake duct 44.

The duct arrangement further comprises an outlet duct 50 for eachengine, which extends from the engine exhaust 25 to an exhaust outlet atthe trailing edge of the transition region 8. Each outlet duct 50transitions from a direction generally parallel to the respective enginelongitudinal axis C at the exhaust nozzle 25, to a direction generallyparallel to the aircraft longitudinal axis A at an exhaust of the outletduct located at the trailing edge of the transition region 8.Consequently, the exhaust duct 50 ensures that air exhausted from theengine 10 is transferred in a direction substantially parallel to theaircraft longitudinal axis A, thereby ensuring optimum propulsiveefficiency. The ducts 44, 50 are located within the aircraft body 2, inthe transition section 8.

Hydraulic and electrical control runs are considered to be safetycritical aircraft components. In order to further reduce theconsequences of an uncontained disc failure, redundant first and secondcontrol runs 54, 56 are provided. In the example shown in FIG. 2,control runs 54, 56 comprising either hydraulic pipes or electricalcables are provided extending from the cabin 5 to a control surface suchas an aileron (not shown) on the wing 4 outboard of the starboard engine10 b. The first control run 54 extends inboard of the engine 10 b fromthe forward cabin 5 to a region downstream of the engine 10 near thetrailing edge of the transition region 8 (i.e. past the rotationalplanes D₃₂-D₄₂ in an inboard side of the engine 10 b), then extendsoutboard toward the aileron. On the other hand, the second control run56 extends outboard of the engine 10 b from the forward cabin 5 directlytoward the aileron (i.e. past the rotational planes D₃₂-D₄₂ on aninboard side of the engine 10 b). Consequently, the control runs 54, 56are segregated either side of the engine 10 b at the point where theypass the burst disc planes, and so they are unlikely to both be damagedby a burst disc event, since debris from a burst disc event tends toemanate from one side of the engine.

FIG. 1 also shows the centre of mass M of the aircraft. As can beenseen, the engines 10 a, 10 b are located in a line B extending generallycoincident with the aircraft centre of mass M. The wings 6 are alsogenerally located extending either side of the aircraft centre of massM, such that the centre of lift (which in this case is provided by boththe wings 6 and fuselage 4) is also located at a longitudinal positiongenerally coincident with the centre of mass M. Consequently, trim dragfor the aircraft is minimised, since the centre of mass M and centre oflift L are close to one another. Due to the small amount of trimrequired, ailerons located at the trailing edge of the swept wings 6 mayprovide sufficient pitch authority (and so acting as elevens), in spiteof the relatively short moment arm. Consequently, a conventional tailsurface may be omitted. Wing tip fences may be provided, which maycomprise actuators to provide yaw control. Consequently, a conventionalvertical tail surface could also be omitted, while maintaininglongitudinal static stability. In view of the relatively close spacingof the engines relative to the centre of mass, adverse yaw in the eventof an engine failure is minimised. Consequently, the minimum controlspeed for the aircraft is adequate, in spite of the lack of aconventional vertical tail.

While the invention has been described in conjunction with the exemplaryembodiments described above, many equivalent modifications andvariations will be apparent to those skilled in the art when given thisdisclosure. Accordingly, the exemplary embodiments of the invention setforth above are considered to be illustrative and not limiting. Variouschanges to the described embodiments may be made without departing fromthe spirit and scope of the invention.

For example, the aircraft could comprise a conventional tube and wingaircraft. The aircraft could comprise a larger number of engines, suchas in a distributed propulsion aircraft. Alternatively, a relativelysmall number of engines (such as two) could be provided, with eachengine providing power for a plurality of propulsors such aselectrically driven propellers, in an arrangement similar to applicant'sco-pending UK patent application EP14192431.6, incorporated herein byreference. In such a case, the propulsors (which would be regarded asbeing safety critical aircraft components) would be located at locationson the fuselage and/or wing such that the planes of rotation of theengines do not intersect with the propulsors. The engines could belocated in different regions, such as at the leading edge of thetransition region, or trailing edges of the wing.

Aspects of any of the embodiments of the invention could be combinedwith aspects of other embodiments, where appropriate.

The invention claimed is:
 1. An aircraft having an aircraft longitudinalaxis and comprising: at least one gas turbine engine having an enginelongitudinal axis and comprising at least one compressor or turbinerotor disc defining a respective rotational plane normal to the enginelongitudinal axis, wherein the engine longitudinal axis is angledrelative to the aircraft longitudinal axis such that a respectiverotational plane of the respective engine is nonintersecting with one ormore of: a further gas turbine engine, a fuel tank, a hydraulic controlrun, an electrical control run, a passenger or crew compartment, or anaircraft propulsor; and at least one duct arrangement comprising anintake duct leading from an air inlet to an intake of a respectiveengine, and an exhaust duct leading from an exhaust of the respectiveengine to an exhaust outlet, wherein the intake duct transitions from adirection generally parallel to the aircraft longitudinal axis at theair inlet, to a direction generally parallel to the respective enginelongitudinal axis at the intake, and the outlet duct transitions from adirection generally parallel to the respective engine longitudinal axisat the exhaust of the respective engine, to a direction generallyparallel to the aircraft longitudinal axis at the exhaust outlet.
 2. Anaircraft according to claim 1, wherein aircraft comprises two or moregas turbine engines and associated duct arrangements arranged in a lineextending generally normally to the aircraft longitudinal axis in thehorizontal plane.
 3. An aircraft according to claim 2, wherein the twoor more gas turbine engines are arranged in a line substantiallycoincident with at least one of an aircraft center of mass and anaircraft center of lift.
 4. An aircraft according to claim 1, whereinthe aircraft comprises one of a blended wing body aircraft and a flyingwing aircraft.
 5. An aircraft according to claim 1, wherein the enginesare located at a trailing edge of the fuselage of the aircraft.
 6. Anaircraft according to claim 1, wherein the engines are mounted at apoint of transition between the fuselage and wings.
 7. An aircraftaccording to claim 1, wherein the engines are located within theaircraft wing or fuselage.
 8. An aircraft in accordance with claim 1,wherein the aircraft comprises at least two redundant control runs forat least one aircraft control surface, wherein one control run extendsalong a first side of an engine, and a further control run extends alonga second side of the respective engine.
 9. An aircraft according toclaim 1, wherein the air inlet of the or each intake duct is located ata leading edge of the point of transition between the fuselage andwings.
 10. An aircraft having an aircraft longitudinal axis andcomprising: at least two gas turbine engines, each having an enginelongitudinal axis and comprising at least one compressor or turbinerotor disc defining a respective rotational plane normal to the enginelongitudinal axis, wherein the engine longitudinal axis is angledrelative to the aircraft longitudinal axis such that a respectiverotational plane of one engine of the at least two gas turbine enginesis nonintersecting with any of: another gas turbine engine of the atleast two gas turbine engines, a fuel tank, a hydraulic control run, anelectrical control run, a passenger or crew compartment, and an aircraftpropulsor; and at least one duct arrangement comprising an intake ductleading from an air inlet to an intake of a respective engine, and anexhaust duct leading from an exhaust of the respective engine to anexhaust outlet, wherein the intake duct transitions from a directiongenerally parallel to the aircraft longitudinal axis at the air inlet,to a direction generally parallel to the respective engine longitudinalaxis at the intake, and the outlet duct transitions from a directiongenerally parallel to the respective engine longitudinal axis at theexhaust of the respective engine, to a direction generally parallel tothe aircraft longitudinal axis at the exhaust outlet.
 11. An aircraftaccording to claim 10, wherein aircraft comprises two or more gasturbine engines and associated duct arrangements arranged in a lineextending generally normally to the aircraft longitudinal axis in thehorizontal plane.
 12. An aircraft according to claim 11, wherein the oneor more gas turbine engines are arranged in a line substantiallycoincident with at least one of an aircraft center of mass and anaircraft center of lift.
 13. An aircraft according to claim 10, whereinthe aircraft comprises one of a blended wing body aircraft and a flyingwing aircraft.
 14. An aircraft according to claim 10, wherein theengines are located at a trailing edge of the fuselage of the aircraft.15. An aircraft according to claim 10, wherein the engines are mountedat a point of transition between the fuselage and wings.
 16. An aircraftaccording to claim 10, wherein the engines are located within theaircraft wing or fuselage.
 17. An aircraft in accordance with claim 10,wherein the aircraft comprises at least two redundant control runs forat least one aircraft control surface, wherein one control run extendsalong a first side of an engine, and a further control run extends alonga second side of the respective engine.
 18. An aircraft according toclaim 10, wherein the air inlet of the intake duct is located at aleading edge of the point of transition between the fuselage and wings.